Stefan Salameh Victoria Mello Aron Potur Boeing Green Cruiser 7G7 Stefan Salameh Victoria Mello Aron Potur

Table of Contents Introduction Background Problem Project Goals Schedule Project Costs Mission Trade Methodology Assumptions Configuration Performance Aerodynamics Propulsion Stability and Control Materials Cost Estimates Conclusion References/ Acknowledgements

Introduction Major challenges in aircraft industry: global warming, impact on environment Global warming: Aircraft blamed for substantial portion, severe restrictions / penalties due to atmospheric pollution levels per passenger. Environmental effects: aircraft manufacturing and disposal Introduction The environment is one of our major challenges today. Global warming will become a big issue in the relatively near future Aircraft may be blamed for a substantial portion of this and severe restrictions or penalties may be placed on aircraft as a function of their atmospheric pollution levels per passenger Another concern is the environmental effects of aircraft manufacturing and disposal

Background “Greener by Design” focuses on the following environment concerns: noise, local air quality and climate change. Continuing retirement of older aircraft from fleet: offsets the effect of traffic growth, noise exposure around airports declining Regulations for aircraft limit pollutants: NOx Measures to reduce fuel burn effect climate change Areas of uncertainty: effects on the ozone created by NOx and the effects of contrails Background The study done in “Greener by Design” focuses on the following environment concerns: noise, local air quality and climate change. According to “Greener by Design:” At present the continuing retirement of older aircraft from the fleet more than offsets the effect of traffic growth, with the result that noise exposure of communities around airports is declining. Regulations for aircraft seek to limit the emission of a range of pollutants, NOx. A reduction in NOx would partially alleviate the impact of increasing air travel not only on air quality local to airports but also on climate change. Measures to reduce fuel burn may significantly increase the overall contribution to climate change. Two particularly important areas of uncertainty are the effects of the ozone created by NOx emissions and the effects of contrails.

Problem “Greener by Design”- design range at which the payload fuel efficiency is a maximum Predicts optimum design range of 2160nm Use “reverse engineering” to establish a baseline transport, compared to the 98,000 lb payload Check to see how max range for max payload efficiency varies with design payload Design an air transport system that is greener overall Problem According to “Greener by Design” there is a design range at which the payload fuel efficiency, defined as the weight of payload times the range divided by the weight of fuel burned, is a maximum. “Greener by Design predicts that the optimum design range is 2160nm We will start by doing some “reverse engineering” to establish a baseline transport design that can be compared to the (98,000lbs) 44.8 tonne payload discussed in “Greener by Design” Next we will check to see how max range for max payload efficiency varies with design payload Finally we will devise an air transport system that is greener by design

Project Goals Perform mission trades Design aircraft to maximize payload fuel efficiency. Perform technology sensitivity trade studies to understand their effect on payload fuel efficiency. * Payload Fuel Efficiency = (Pounds of payload X range) / pounds of fuel

Schedule

Project Costs Time: 10 to 12 hrs per week, per person. Therefore, approximately 100 to 120 hrs x 3 engineers, so roughly around 300 to 360 hours depending on the design needs, and the time available around our own schedules. Using the RE of $95.6 (accounting for the 2004 CPI) approximate engineering value of this design is $28,680 to $34,416. One or two trips to Boeing. One vehicle, approximately 20 mile per gal, about 100 miles, $2.30 per gallon, this brings the cost of the trip(s) to $11.5 or $23 (two trips).

Mission Trade Methodology Fuel Fraction Buildup Method L/D and TSFC estimated from historical trend data. Perform sensitivity studies To size the aircraft we used the fuel fraction method to find the fuel burned by each segment. We used the Breguet range equation where the main variables are TSFC and L/D. Our speed and altitude are constant at 36000 feet and Mach 0.8, L/D and TSFC were varied.

Assumptions (2020 predictions)

TSFC Use of historical data to estimate the TSFC for 2020 Estimated 2020 TSFC of 0.45 NASA’s Ultra Efficient Engine Program (UEEP) has made this plot of historical trends for engines. Their intent is to look at the future use of Very High Bypass ratios for turbofan engines. The engine sometimes has a name “ducted fan”. Since we believe aircraft manufacturers will step towards higher bypass ratios on turbofan engines, we use one in our design. The unducted fan may be the successor to the very high bypass turbofan, but its development is stalled due to complications. Use of historical data to estimate the TSFC for 2020 NASA’s Ultra Efficient Engine Technology Program

L/D This plot of L/D max trends over the course of the past 50 years shows a slight improvement over time. There is, however not much of an increase. The higher end of these values hovers around 18. This is a value that can be reached with today’s drag reduction methods. The best range L/D would be around 15.5. This number will be used in studies of payload fuel efficiency.

Estimating MTOGW & OEW Compiled current data Predicted 8% saving in weight Material advances will allow for lighter OEW for a given MTOGW The method used was detailed in Schaufelle’s aircraft design book. The method consisted of fuel fraction build up. The mission profile is shown below. All segments besides the cruise portion had standard fuel fractions in Schaufelle. The cruise fuel fraction comes directly from manipulation of the Breguet range equation.

Configuration

The 3–View

Top View Interior Layout Aircraft can be configured for more seats and two class. This will increase the pax weight and revenue cargo will be reduced. Note does not include fully sized tail at the left end.

Design Layout

Coach 224 Seats Pitch = 32 inches Partitioned Into Two Sections With 8 Rows Each As you see, the cabin is essentially two overlapping fuselages. It must be noted, due do pressurization considerations, there must be some support added in the center of slightly off center of the fuselage. Support members could be spaced to be hidden in the bulkheads between cabins.

Business 58 Seats Pitch = 40 inches Private Lavatories and Galley Not much to say here.

First Class 14 Seats Pitch = 60 inches Private Lavatories and Galley The bar in the first class cabin is one of the layout option that could be incorporated into the design.

Emergency Egression For the paz capacity of 298 passengers, the aircraft is required to have 10 emergency exits per FAR 25.807 requirements. The flowing exits are included. 3-Type A floor level exits each side 2-Type III over wing exits each side. The over wing exit is two side by side

Performance

Range . Here is the payload fuel efficiency for various payloads. It is interesting to see that payload does not play a big factor on best range for fuel efficiency. The values used for this study our L/D of 18 and a TSFC of .45. These are estimates we believe are attainable by 2020. The greater the payload, the better the payload fuel efficiency.

Payload If we assume from the last page optimal range is around 2860 and plot payload fuel efficiency for a range of payloads, we see that there is a theoretical limit to payload fuel efficiency. This leads away from the notion that an infinitely large aircraft would be the way to go.

Payload Taking a closer look at more reasonable payloads, we see an area where a change in payload does not cause a big change in payload fuel efficeincy. This is desirable if the aircraft is to be under loaded and flown slightly longer range or vice versa. The upper end of our design box represents a limit to payload that we believe will keep structural weights from becoming to large. Since the aircrafts will not have a range like most long range wide-bodies, it will carry much less fuel while still holding the same amount of payloads in the fuselage. The bending moment of the wings will not have as much assistance from fuel in the wings. To keep wing weights from being high, the max payload should be under 110,000. Also, past 110,000 lbs payload, there is not much increased benefit of payload fuel efficiency for a rise in payload.

Choosing Payload Payload fuel efficiency reaches theoretical asymptote of 3750 lbs*nm/lbs Payloads within 10% to 15% of maximum payload fuel efficiency seem to show a “knee” in the curve. We will focus on a design box with a payload of 90k lbs. The previous slide explains some of the reasoning.

Effects of L/D and TSFC No significant effect with varying L/D from 15 to 25 on “optimal” range. TSFC does play important role. As TSFC decreases, optimal range increases. We will stay with .45 1/hr as our TSFC. Increasing L/D does increase payload fuel efficiency for a given payload. It does not, however, influence the best range that the aircraft is suited for. Changing TSFC has a big effect on the optimum range.

Takeoff / Landing Distances Takeoff Distance = 4268 feet Landing Distance = 3717 feet

Aerodynamics

Summaries

Summaries (cont.)

Airfoil Selection Wing Root Wing Tip Here we see the use of a critical airfoil for the 7G7. This is just a generic airfoil . The wing hub has a thickness ratio of .15 and tapers to the tip which has a t/c of .1. The plots were made through VisualFoil. The wing hub has transition to turbulent flow at 15%. The tip is able to delay transition to 50% on the lower surface.

Laminar Flow Research From the Lockheed Martin team, we were able to use some data on passive laminar flow control. It is clear to see that a t/c of .15 on this specific airfoil (CAST 10-2/DOA 2) leads to laminar flow over 60% of the wing. Our design incorporated a critical airfoil, which while good for transonic drag reduction, may not be the best choice to delay boundary layer transition.

Wing Planform Ref. Area = 3218 Aspect Ratio = 7.0 Sweep = 30 Taper = 0.2036 Span = 150.1 Length = 57.67 Gross Area = 3458.32 Exposed Root = 41.43 M.A.C. = 27.92

Lift Distribution This planform shown is representative our final selection. Using the airfoil in the previous slides, this is the wing loading distribution during cruise conditions. This data could be further used by a structures group in estimating a revised wing weight.

Propulsion

Engine Selection and Placement Very High Bypass Turbofan Engine Large Fan Size Makes Engine Dimensions Similar to GE-90 Due to the high bypass ratios, the fan size of the engine is similar to the GE-90 engine, that used in the 777. The aircraft’s MTOGW is half the 777 and the over wing structure is smaller. To keep the aircraft from needing a large landing gear structure, the engines are mounted on the rear of the aircraft. The need for no engine pylons on the wings reduces the interference drag and may aid in keeping the flow laminar further down the wing.

Engine Characteristics The engine was sized with methods discussed in Schaufele

Engine Geometry / Sizing

Stability and Control

Stability and Control T-Tail incorporated to compliment the rear engine layout. Neutral point located close to CG will allow for smaller horizontal stabilizer Smaller vertical stabilizer needed due to less single engine yaw effects. SAS used for longitudinal and directional stability. Estimations for the CHT and CVT were taken from table 6.4 of Raymer. CVT = 0.0855 * ((0.09 – (0.09*0.05)) CHT = 0.95 * ((1.00 – (1.00*0.05))

T-Tail Selection / Sizing LVT = 76.3 ft LHT = 81.4 ft SVT = 478.6 ft² SHT = 811.2 ft² Vertical tail volume coefficient can be reduced by 5% due to end-plate effect. Similarly the horizontal tail volume coefficient can also be reduced by 5% due to the clean air seen by the horizontal tail. (Raymer p.125) In order to estimate the moment arm of the vertical tail (LVT), which is defined as the distance from the quarter-chord of wing to the quarter-chord of vertical tail, and the moment arm of the horizontal tail (LHT), which is defined as defined as the distance from the quarter-chord of wing to the quarter-chord of horizontal tail we first need to the locations of the quarter chord of the wing, vertical and horizontal tails. Since the fuselage length is 2035 inches (169.6 feet), Raymer’s design guides indicated that the moment arm of the T-tail, of both the vertical and horizontal tail be about 45% - 50% of the fuselage length

Longitudinal, Lateral, & Directional Control Surfaces Outboard Aileron (low speed) Inboard Aileron-Flaperon (High speed bank control and flapped for takeoff and landing.) Inboard Flaps – Double Fowler Outboard Flaps – Double Fowler Single Rudder One Elevator per side of horizontal stabilizer Single Slat on Leading Edge The model of the aircraft was completed in SolidWorks. Control surfaces consist of an outboard aileron, outboard flap, inboard flaperon, and main inboard flap. Flaps are tripple slotted fowler flaps. A one piece slat extends across a majority of the leading edge and can be heated with bleed air or by electric means (perhaps piezoelectric means in a composite material). The vertical stabilizer has a hinged rudder of one single piece and the horizontal stabilizer has two elevators that work in unison.

Weights / Balances

CG Travel Using basic methods discussed in Raymers text, the mass and cg of the various components was estimated.

Materials 2020 Reductions in weight of the airframe: lighter materials, advances in structural design Conventional aluminum alloys and steels replaced by carbon-fiber reinforced plastics (CFRP) and titanium alloys Advantages: reduced costs of fuel, reduction in fuel burn, replacement of damaged CFRP with aluminum Disadvantages: extra capital cost of the aircraft, cost of repairs to CFRP components expensive Materials Reductions in weight of the airframe by use of lighter materials and advances in structural design Conventional aluminum alloys and steels have been replaced by carbon-fiber reinforced plastics (CFRP) and titanium alloys Advantages: reduced fuel costs arise from the weight savings, reduction in fuel burn, cost of fuel, possible replace damaged CFRP components with aluminum replacements Disadvantage: extra capital cost of the aircraft with the lighter material, cost of repairs to CFRP components being approximately twice that of aluminum alloy components

Materials 2020 Charts taken from Greener by Design Materials weight distribution predictions : conservative forecast

Materials 2020 Charts from Greener by Design Materials Weight Distribution predictions: optimistic forecast

Landing Gear for 2020 Supplier: Messier-Dowty Centerline Main Landing Gear Swept inwards Four wheel bogie design 2 under wings Main wheels 46x16 R20 in Shortening mechanism and articulating bogie Nose Landing Gear Forward Retracting Twin wheel design Nose wheels 40x14 R16 in Titanium main fitting Introduction of 350-bar hydraulic pressure Self explanatory

Landing Gear 2020 New coatings: use of HVOF (High Velocity Oxygen Fuel) surface treatment process instead of using Chromium VI plating Cost Reductions -To minimize lead time, time to manufacture and maintenance Weight Reductions -Titanium main fitting Noise Reductions -noise reduction add-on devices -installation of fairings around main landing gear. The reductions for cost, weight and noise were based studies and on predictions from Messier-Dowty and Airbus Industries.

Cost Estimates Engineering hours = HE = 7.07 * We0.777 * V0.894 * Q0.163 Tooling hours = HT = 8.71 * We0.777 * V0.696 * Q0.263 Manufacturing hours = HM = 10.72 * We0.82 * V0.484 * Q0.641 Developing support cost = CD = 66.0 * We0.630 *V1.3 Flight test cost = CF = 1807.1 * We0.325 *V0.822 * FTA1.21 Manufacturing materials cost = CM = 16 * We0.921 *V0.621 * Q0.799 Engineering production cost = Ceng = 2251* [0.043 * Tmax + 243.25 *Mmax + 0.969 * Tturbine inlet – 2228] RDT&E + flyaway = HERE + HTRT + HMRM + HQRQ + CD + CF + CM + CengNeng + Cavionics CPI information was found online from http://www.bls.gov/ U.S. Department of Labor, Bureau of Labor Statistics and then it was verified using an online U.S. Department of Labor calculator (http://data.bls.gov/cgi-bin/cpicalc.pl). The formula used for the CPI compensation: 2004 Price = 1999 Price x (2004 CPI / 1999 CPI) Once the aircraft has been conceptually designed & analyzed and then submitted for proposal the customer will have a number of aircraft that all meet the requirements. Naturally they will all differ in design and thus differ in their performance, but again they all meet the requirements. This is where the customer will definitely take notice of the cost of the aircraft.

Cost Estimates It is estimated that the RDT&E and Production costs for this program are around $192 million per aircraft. The total Acquisition cost for all 100 aircraft equals $19.2 billion. RDT&E (Research, Development, Test, and Evaluation) includes all the technology research, design engineering, testing, and other evaluations, within it is the cost of the conceptual design. RDT&E costs are fixed and don’t depend on how many aircraft are produced; they are usually less than 10% of the total life cycle cost. Cost estimation data is for a number of aircraft is used and analyzed to make “cost estimating relationships” (CER). The Development and Procurement Costs of Aircraft model (DAPCA) is a set of CERs for a conceptual aircraft design. It provides a reasonable cost estimate for a number of different aircraft. The equations given can be calculated in the meter-kilogram-second system (mks) or the foot-pound-second system (fps), for our purposes and calculations we chose the foot-pound-second system. RDT&E + flyaway cost calculated above is for the 100 aircraft and the value obtained is 19,196,065,759. Therefore per aircraft RDT&E + flyaway is 191,960,657 or just under 192 million constant 2004 US dollars. Raymer provides “fudge factors” for the hours of the DAPCA equations in order to correct for the difficulty in constructing an aircraft with complex materials such as fiberglass, graphite-epoxy, titanium and so on. Since our aircraft is proposed to be built with 2020 technology, we have used Raymer’s fudge factors to obtain the above values. INSERT GREEN IMAGE FROM VICTORIA’S FILE

Payload Fuel Efficiency Comparisons Here is a comparison of many current transport aircraft. The data used to compute the Payload Fuel Efficiencies for the various aircraft is done by using the design range, design payload, and design fuel weight. The data was taken from data sets provided by Civil Jet Aircraft Design by Butterworth-Heinemann. And the 7G7 wins!

References & Acknowledgements Green, J.E. “Greener By Design,” Aeronautical Journal vol. 106, no 1056 Feb. 2002 Schaufele, Roger D. “The Elements of Aircraft Preliminary Design,” Aries Publications 2000 Raymer, Daniel P. “Aircraft Design: A Conceptual Approach” 3rd ed. AIAA 1999 www.bh.com/companions/034074152X/appendices/default.htm